The present invention relates generally to gas turbine engines, and, more specifically, to turbines therein.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases that flow downstream through a high pressure turbine nozzle which directs the flow into a row of high pressure turbine rotor blades. The blades extract energy from the gases for powering the compressor, and a low pressure turbine is disposed downstream therefrom for extracting additional energy which typically powers a fan for producing propulsion thrust to power an aircraft in flight.
The nozzle includes stator vanes which like the turbine blades have an airfoil or crescent profile with substantial curvature or camber between leading and trailing edges thereof. The vane and blade airfoils have a generally concave pressure side and an opposite generally convex suction side along which the combustion gases flow during operation.
The respective suction sides of the vanes and blades are circumferentially spaced from the pressure sides of adjacent vanes and blades to define corresponding flow channels therebetween through which the combustion gases flow. The combustion gases enter the turbine nozzle in a generally axial downstream direction and are redirected at the trailing edges of the vanes at an oblique angle toward the leading edges of the rotating turbine blades.
Accordingly, the individual streamlines of the combustion gases flow generally parallel to each other between the nozzle vanes and between the turbine blades, but vary in curvature to correspond with the different velocities thereof as effected by the suction and pressure sides of the adjacent vanes and blades.
The blade platforms define the radially inner boundary which bounds the combustion gases as they flow between the turbine blades, and those platforms are aerodynamically smooth for maximizing efficiency and performance of the turbine during operation. However, the blades are individually mounted in the perimeter of a corresponding rotor disk using corresponding integral blade dovetails. And, each blade includes an individual platform integral with the airfoil and dovetail thereof, which platforms must circumferentially adjoin each other with a minimum circumferential clearance or gap therebetween.
Since the blade platforms are subject to common manufacturing tolerances which randomly vary their final dimensions, and since the platforms are also subject to stackup tolerances when mounted in their corresponding dovetail slots in the disk perimeter, the adjacent side edges of the platforms which define the axially extending gap therebetween are subject to random differences in radial elevation or height.
Should those platform side edges create an upstream facing step, the downstream flowing combustion gases will impinge on the up-step resulting in loss of aerodynamic efficiency as well as local heating of the projecting step leading to early oxidation and locally high thermal stress which may reduce the useful life of the turbine blade.
Preferably, the platform side edges should be flush or have a slight downstream facing step over which the combustion gases may flow without obstruction. However, since the combustion gas streamlines are necessarily redirected from the nozzle vanes between the turbine blades, those streamlines will normally traverse the platform gaps in opposite circumferential directions at the inlet and outlet ends of the corresponding flow channels.
This adds to the complexity of the platform design which may be intentionally warped to create the down step in one circumferential direction at the inlets to the turbine blades, and another down step in an opposite circumferential direction at the outlets of the turbine blades.
In order to create these oppositely facing down steps, the suction-side edge of one platform must be radially higher than the pressure-side edge of the adjacent platform at the forward ends thereof, with the aft ends thereof being radially lower for the suction-side edge compared with the pressure-side edge of the adjacent platforms. In this way, the steps between adjacent platforms change magnitude and direction between the forward and aft ends of the platforms for ensuring down steps only relative to the local direction of the combustion gas streamlines.
However, since the inter-platform steps change direction in this configuration, the steps gradually change in magnitude between the forward and aft edges of the platform and transition through a no-step flush point of the platforms in the axial middle thereof. Since the local direction of the combustion gas streamlines can change over the different operating points of the engine as well as over the life of the engine, the platform steps at both ends of the transition point may still create undesirable up steps instead of the intended down steps. An up step between the platforms locally obstructs the smooth flow of the combustion gases, and is locally heated thereby with an increased heat transfer coefficient, and is therefore subject to increased oxidation and locally high thermal stress which may decrease the useful life of the turbine blades.
Accordingly, it is desired to provide an improved turbine blade with a step-down platform for substantially all operating conditions of the gas turbine engine.
A turbine includes a row of blades each having an integral airfoil, platform, and dovetail. Each platform has opposite first and second side edges corresponding with the opposite pressure and suction sides of the airfoil. The platform first side edge is radially higher than the platform second side edge continuously between opposite forward and aft edges of the platform. In this way, adjacent platforms define a down step therebetween for preventing obstruction of combustion gases flowable downstream thereover.